United Technologies Corporation

Alstom Technology Limited

Other

Power Systems Mfg., LLC

Ansaldo Energia IP UK Limited

Douglas James Dietrich

Stephen Wayne Fiebiger

Gregory Edwin Vogel

Alan W Stoner

Adam Lee Hart

Frank W Huber

Alex Torkaman

W David Day

Kevin Damian Carpenter

Alexander Vasili Rozendaal

Christopher Charles Williams

Richard Ramirez

Stephen Fiebiger

David J Wiebe

Kuo-Ting Hsia

Samer Abdel-Wahab

Adam John Fredmonski

Lonnie Adolphus Houck, Iii

Roy E Struble

Megan Burke

Jeff Tessier

A gas turbine blade for a gas turbine engine includes an attachment, a blade neck portion, a platform, an airfoil, and a shroud. The gas turbine engine has an operating speed range. The blade neck portion extends radially outward from the attachment. The platform is attached to the blade neck portion opposite the attachment. The airfoil extends radially outward from the platform. The airfoil includes a leading edge, a trailing edge, and an axial chord defined between the leading and trailing edges. The axial chord defines a mid-chord point. The shroud is attached to an end of the airfoil. The shroud includes an upper surface and a modal response tuning component attached to the upper surface of the shroud. The modal response tuning component is offset from the mid-chord point a predetermined distance to move modal response of the gas turbine blade outside of the operating speed range.

Turbine blade with modal response adapted tip shroud

A turbine blade, airfoil, and rotor stage for a gas turbine engine is disclosed. The turbine blade, airfoil, and rotor stage each includes an uncoated airfoil profile in accordance with Cartesian coordinate values of X, Y, and Z disclosed herein. The resulting airfoil exhibits high flutter margins, thus enabling the gas turbine engine to be operated at an increased power output while avoiding operational limitations required in known gas turbine engines.

Turbine blade having improved flutter capability and increased turbine stage output

The present invention relates to a system and method for joining adjacent sections of a compressor diaphragm. A splitline jumper is positioned within a recessed portion of a seal ring segment of the adjacent compressor diaphragms. The splitline jumper is sized and oriented so as to prevent axial movement of the compressor diaphragms relative to each other and also to prevent rotational movement of the seal ring relative to the diaphragm.

Anti-rotation lug and splitline jumper

A turbine blade, airfoil configuration, and rotor stage are disclosed in which through the airfoil profile disclosed in Table 1, a modification in airfoil flutter and swirl are achieved. Through the airfoil configuration, the swirl and improved platform sealing configuration, result in the turbine blades having the airfoil profile with an increased performance output from the turbine.

A system and method of extending the useable life of a gas turbine blade is disclosed in which the gas turbine blade includes an undercut configuration designed to relieve mechanical and thermal stress imparted into the pedestal region of the airfoil trailing edge. The embodiments of the present invention include turbine blade configurations having different trailing edge undercut configurations as well as additional cooling supplied to the internal passages of the trailing edge region of the turbine blade.

Turbine blade platform undercut

The internally air cooled turbine cast blade is molded with a ceramic core for defining the interior cavity with coolant passages where the ceramic core is configured to fair radially inwardly at the attachment portion of the blade until they form a single wall, parallel to and centered in the blade attachment. The blade is cast in the mold by pouring molten metal into the mold, curing and the core is leached once the metal is cured and the mold removed. Another embodiment includes a metering plate attached to the root of the blade for metering coolant to provide a coolant to gas flow ratio at the film cooling holes.

Method of configuring a ceramic core for casting a turbine blade

Turbine blade construction

Ceramic core for casting a turbine blade

A device for sealing between and damping of adjacent rotor blade platforms is provided which includes a forward seal member, a compliant center seal member, an aft seal member, a first damper, and a second damper. The first damper is disposed between the forward and compliant center seal members, and the second damper is disposed between the compliant center and aft seal members. The compliant center seal member is disposed between the first and second dampers. The compliant center seal member accommodates misalignment between adjacent rotor blade platforms by deflecting in response to centrifugal forces acting on the dampers.

Turbine engine rotor blade platform sealing and vibration damping device

An aerodynamic buffer zone in the gap between the tip of an air cooled axial flow turbine blade of a gas turbine engine and its surrounding shroud or seal (outer air seal) is created by discretely discharging a high velocity stream of cooling air from the tip of the blade in a judicious direction. The stream is angularly oriented relative to the plane of rotation of the turbine so that it is discharged in the gap in the direction of the pressure side of the turbine and preferably at between a 15.degree.-45.degree. angle relative to the flat surface of the tip of the blade.

Clearance control for the turbine of a gas turbine engine

Growing community of inventors

West Palm Beach, FL, United States of America

Douglas James Dietrich